The field of the disclosure relates generally to gas turbine engines and, more particularly, to a method and system for controlling compressor clearance at various stages of flight using active cooling of the compressor case.
Gas turbine engines typically include multiple compressor stages to compress incoming air flow for delivery to the combustor. The rotor blades and compressor casing are subjected to a range of temperatures during various stages of operation such as ground operation, takeoff, and cruise, resulting in thermal expansion or contraction of these compressor components. Typically, the components of the compressor stages are designed to operate with minimal rotor tip clearances and interstage seal clearances to enhance thrust production during takeoff. However, during cruise conditions, operating temperatures of the compressor stages are lower than at takeoff, resulting in higher clearances due to thermal contraction of the compressor components. Higher rotor tip and interstage seal clearances degrade the efficiency of operation of the gas turbine engine at cruise conditions. A reduction in rotor tip and interstage seal clearances at cruise conditions, without impacting the operation of the gas turbine engine at takeoff conditions, can enhance fuel efficiency of the gas turbine engine during cruise conditions with minimal impact on thrust production at takeoff conditions.